Beam capture missile guidance system



1967 D. J. CAMPBELL ETAL 3,356,314

BEAM CAPTURE MISSILE GUIDANCZ SYSTEM Filed Oct. 24, 1963 4 Sheets-Sheet l TARGET DATA sIIIPs TURN XMITTER ANEMOMETER XMITTER 232 A A GYRO COMPASS STABILIZATION 1 COMPUTER r TRACKING A RADAR GUIDANCE I III RADAR 22 16 c 1 I o M K WEAPON SEARCH A iE g 11 k GUIDANCE I m RADAR 21 TRACKING A RADAR STABILIZATION 4 COMPUTER LAUNCHING SYSTEM ,INVENTORS TYPE B DAVID J. CAMPBELL MELVINQLEE F/g 2 BY JEROLD F. MANN SURFACE OF EARTH L ATTORNEY 1967 D. J. CAMPBELL ETAL 3,356,314

BEAM CAPTURE MISSILE GUIDANCE SYSTEM Filed Oct. 24, 1963 4 Sheets-Sheet 2 K(R-Rm) RATE ORDER FROM COMPUTER A S e 61 62 lf f GYRO PREAMP 4 I l l 63 I I l 1 AMP l l I l i 64 I 66 I INVENTORS RATE SENSOR DAVID J. CAMPBELL BY MELVINQLEE PO$|T|QNAL JEROLD F. MANN y 6 sHIPs MIOTION OUTPUT ATTORNEY 1967 D. J. CAMPBELL ETAL 3,

BEAM CAPTURE MISSILE GUIDANCE SYSTEI" Filed Oct. 24, 1963 4 Sheets-Sheet .3

BORESIGHT DATUM PLANE B E A NL AXIS BORESIGHT TARGET L08 DATUM PLANE BORESIGHT DATUM PLANE MISSILE LOS INVENTORS DAVID J. CAMPBELL MELVINQLEE BY JEROLD F. MANN TARGET LOS ATTORNEY 4 Sheets-Sheet 4 GUIDANCE RADAR fig. i@

INVENTORS DAVID .1. CAMPBELL ATTORNEY MELVINQLEE JEROLD F. MANN CROSS TRAVERSE RATE CIRCUIT D. J. CAMPBELL ETAL BEAM CAPTURE MISSILE GUIDANCE SYSTEM INPUT COMPUTER GUIDANCE RADAR SERVO Dec. 5, 1967 Filed Oct. 24, 1963 FUNCTION GENERATOR TYPE B TRAJECTORY United States Patent C 3,356,314 BEAM CAPTURE MISSILE GUIDANCE SYSTEM David J. Campbell, East Meadow, Melvin C. Lee, Westbury, and Jerold F. Mann, Port Washington, N.Y., as-

signors, by mesne assignments, to the United States of America as represented by the Secretary of the Navy Filed Oct. 24, 1963, Ser. No. 319,094 13 Claims. (Cl. 2443.13)

This invention relates to a computer system and more particularly to a computer system for programming and transmitting guidance information to a beam riding missile.

In a relatively new field of missile guidance, various techniques are used to guide a missile to a predetermined destination. The guiding systems for long range missiles are usually highly complicated inertial guidance systems utilizing computer techniques based on earth or space references. In these missile guidance systems the missile itself contains all or most of the guidance equipment necessary for its own guidance. Thus, the missile guidance system is self contained in a missile and once destination information is set into the guidance system, the missile in effect guides itself to the predetermined destination.

When anticipated targets are within radar range, it is possible to guide a missile to a predetermined destination by external beam guidance techniques. Thus, the bulk of the guidance system may be external to the missile and the missile contains a minimum amount of guidance equipment.

The present invention contemplates a computer for providing mid-course guidance for short range anti-aircraft missiles. Mid-course guidance begins when the missile is captured in the programmed guidance beam and ends when homing starts for the self-homing type missile or at target intercept for the remotely detonated missile. After launching of the missile, the missile is captured by the radar guidance beam and thereafter the missile operates as a beam rider. This invention provides a guidance beam program which has a high degree of space stabilization from capture to target intercept. Since excessive residual beam motion or jitter induces unnecessary drag on the missile, the present invention contemplates a beam guidance which substantially eliminates residual beam motion or jitter.

During initial guidance the computer generates outputs which position the guidance beam to cause the missile to follow an up and over trajectory in one mode while in another mode the guidance beam causes the missile to follow a line of sight path to the target. Interception of the missile with the target is effected by bringing the missile and target tracking beams together as the missile reaches the target, i.e., as a function of change in range between target and missile. On arriving at the target the missile warhead may be detonated by a proximity fuse at the end of a homing program initiated somewhat previous to target contact. Alternately beam riding may be continued until there is range coincidence at which time detonation may be commanded remotely.

The missile guidance programmer computer of the present invention forms the principal guidance system of a more complex weapons control system. The weapons control system which forms no part of this invention selects, tracks a target, and calculates all necessary data for launching the missile into the guidance radar beam of the present invention. After the missile is launched and captured by the beam, it is directed to the target by the guidance information contained in the radar guidance beam provided by the present invention. During the course of missile flight the guidance information is furnished to the guidance radar set by the computer based on the continu- 3,356,314 Patented Dec. 5, 1967 ous tracking information being received from a tracking radar. Upon reaching the vicinity of the target, the missile receives tracking radar energy reflected from the target which it utilizes to home on the target and subsequently to be detonated by a proximity fuse. Alternately the missile may ride the guidance beam all the way to the target and be detonated at the proper instant by a commanded signal transmitted in the guidance beam.

Therefore, it is an object of the present invention to provide a computer system for programming and transmitting guidance information to a beam riding missile;

Another object of the present invention is to provide a computer which selectively programs a missile flight in one of two modes by transmitting a modulated guidance beam to a beam riding missile for guidance of the missile to target intercept.

A further object of the present invention is to provide a mid-course guidance program for a beam riding missile which has a high degree of space stabilization to effectively reduce adverse drag on the missile caused by residual beam motion or jitter.

With these and other objects in view, as will hereinafter more fully appear, and which will be more particularly pointed out in the appended claims reference is now made to the following description taken in connection with the accompanying drawings in which:

FIG. 1 illustrates the weapons control system of which the present invention forms a part;

FIG. 2 illustrates the two missile trajectories which may be programmed by the present invention;

FIG. 3 is a graphical illustration of the type A beam program;

FIG. 4 is a graphical illustration of the type B beam program;

FIG. 5 illustrates a second-order smoothing circuit utilized by the present invention;

FIG. 6 is a simplified block diagram of the guidance transmitter servo loop;

FIG. 7 is a vector diagram for the type A program;

FIG. 8 is a vector diagram for the type B program;

FIG. 9 illustrates the coordinate transformation required for component development; and

FIG. 10 is a more detailed block diagram of a portion of the system of FIG. 1.

Referring now more particularly to FIG. 1 there is shown the weapons control system of which the present invention forms an integral part and which is disclosed to set forth the environment of the present invention to provide an adequate understanding and appreciation of the present invention. The weapons control system as shown in FIG. 1 is an integrated missile guidance system primarily designed as an anti-aircraft device which functions to select or designate, track, intercept, and destroy high speed enemy aircraft. The particular type of missile for which the weapons control system was developed is a short range ram jet propelled beam-rider type whose warhead may be detonated upon receipt of a command signal from the fire control system when the missile is within destruction range of an enemy aircraft or whose warhead may be detonated by a proximity fuse automatically set off when the missile is within destruction range of the enemy aircraft. 1

The weapon control system shown in FIG. 1 comprises the following major components: the general dual simplex computer system 11 of which Sections III form the subject of the present invention, two radar tracking sets 12 and 13, two radar guidance sets 14 and 15, one weapon direction system 16, one target data transmitter 17, one ships turn transmitter 18 and two stabilization computers 19 and 21. The output of a search radar 22 is connected directly to the weapon direction system 16.

The general dual simplex system 11 receives specific target assignments and technical commands from the weapon direction system 16 which has outputs connected individually to Section I of general dual simplex computer 11. Target positional data from search radar 22 and pitch and roll from gyro compass 23 are fed to the general dual simplex computer 11 to initiate the target designation phase. Own ships data is obtained from ships turn transmitter 18, stabilization computers 19 and 21, anemometer 24, and are fed as shown into general dual simplex computer 11. The general dual simplex computer system 11 is designed to handle two targets simultaneously. It is therefore a dual channel device which accounts for the duplication of tracking guidance radars, etc.

'Where considering the case of one target, a tracking radar, for example tracking radar 12, acquires target position data and on this basis tracks the target automatically providing continuous outputs to the general dual simplex computer 11 which smooths this positional information, generates rates and provides a prediction of intercept position. The general dual simplex computer 11 develops necessary computations and corrections of the intercept data for the systems use and transmits orders based on these data to launching system 26 and to guidance radar 14. Where there are two targets, tracking radar 13 and guidance radar 15 will also be functioning.

The orders from Section II of general dual simplex computer 11 position the launching'platform and the guidance radar 14 so that the missile will be launched into the guidance radar beam. When the missile is launched and captured by the guidance radar beam it is directed to the target by the guidance radar beam from guidance radar 14. Throughout the course of its flight, guidance information is furnished to the guidance radar 14 from general dual simplex computer 11 which information is based on the continuous tracking information being received from tracking radar 12. The missile uses this information to home on the target and subsequently to be detonated by a proximity fuse or alternately the missile may ride the guidance beam all the way to the target to be detonated at the proper instant by a commanded signal in the guidance radar beam intelligence.

The above description of FIG. 1 is intended as background information only and no attempt will be made to describe FIG. 1 in rigorous detail since it forms no part of the present invention.

The general dual simplex computer 11 comprises three sections which perform particular functions. Section I of which two are required, performs the:

(a) computation of search orders for radar designation (acquisition and designate mode),

(b) conversion of target coordinate (gimbal computer),

(c) prediction of intercept point and position (prediction computer).

Section II of which only one is required performs the:

(a) generation of capture angles (capture computer),

dead zone computation (dead zone computer),

(b) generation of missile launching orders (launching computer).

Section III which embodies the present invention and of which two are required comprises the guidance control.

Thus, Sections I and II of general dual simplex computer 11 contain preengagement computation equipment control system of FIG. 1 designates target to tracking radars by supplying to the computer East-West and North- South slant range components of target position and the minimum and maximum target height or true target elevation angles. Section I of the computer then generates an elevation search program which causes the tracking radar to scan up and down in true elevation at the designated true bearing.

The maximum and minimum elevations for the scan functions are determined by the designated target height limits and the slant range of the target. The elevation scan rate is an inverse function of slant range so that the probable acquisition times remain approximately the same at all ranges. Smaller acquisition times result whenever the spread between the height limits is decreased or when more accurate height information is available to the weapon control station.

By appropriate means true bearing and elevation angles are transformed into deck angles of the station. These angles are then transmitted as point orders to the tracking radar 12. The predictor circuit contained in Section I of the computer utilizes data representing target present position, intercept position and time-to-intercept to predict a value of time to fire to the computer.

The point orders necessary for missile launching are generated in Section II of the general dual simplex computer 11 and fed to launching system 26. Section II of the general dual simplex computer 11 contains a capture computer unit which generates in stable coordinates the bearing and elevation angles at which the missile will be captured in the guidance beam. At medium and long ranges the capture bearing angle is given by the present position of the target plus a lead angle based on the horizontal velocity and a fixed value for the flight time, thereby resulting in a decreasing lead angle as the range increases. The elevation angles at the ranges are functions of predicted target intercept position such that the maximum allowable superelevation is obtained. At short ranges both bearing and elevation capture angles are directly representative of the intercept position of the target.

Section III of the general dual simplex computer 11 forms the subject matter of this invention. It is this device which programs the missile flight. The beam program generated by Section III is designed to combine the required smoothing and programming functions in a circuit the time constant of which is a function of the remaining time of flight of the missile. The beam programming function of Section III of the general dual simplex computer consists of five basic blocks. These five basic blocks comprise input computer 31, beam sensitivity computer 32, function generator 33, cross traverse rate circuit 34, and smoothing circuit 35.

FIG. 10 illustrates Section III of the general dual simplex computer 11 showing the details of beam sensitivity computer 32 and smoothing circuit 35. Input computer 31, function generator 33 and cross traverse rate circuit 34 are shown as blocks.

Input computer 31 transforms launcher angles into quantities required by the guidance control system and provides these transformed launch angles as inputs to the smoothing circuit 35 for generating transmitter orders for missile capture. During lift-ofl? of the missile from the launcher, the transmitter beam is held stationary in space. Following capture of the missile by the transmitter beam, input computer 31 provides target angles through smoothing circuit 35 and radar tracking information is processed to produce the appropriate input signals to the smoothing circuit 35 for the guidance mode. All appropriate switching which places the system in aiming memory and guiding modes of operation is included within the input computer 31.

i The beam sensitivity computer 32 which is connected to smoothing circuit 35 varies the sensitivity of the smoothing circuit 35 as a functionof time to intercept. Thus,

sifiooth beam riding characteristics are obtained in the early part of mid-course guidance when the input data to the guidance system may be noisy without sacrificing accuracy when the missile approaches the intercept point. Depending upon tactical conditions, the smoothing circuit may be used to generate alternate trajectories, such as the type A trajectory, or the line of sight trajectory B.

The up and over trajectory type A is realized by a function generated in function generator 33 which when added to the inputs of the smoothing circuit 35 causes the missile to follow an up and over trajectory.

The two trajectories type A and type B are illustrated in FIG. 2. With no input to the smoothing circuit 35 from function generator 33, type B trajectory will be automatically programmed. When function generator 33 is connected to smoothing circuit 35 as more fully discussed hereinbelow, the type A or up and over trajectory is programmed.

Since there may be a dynamic error in the torques applied to the elevation and traverse gyros which position the beam when there is any angular velocity about the line of sight when other angular velocities are present, a cross traverse ra'te circuit 34 is provided to generate the appropriate correction torques for application to the traverse and elevation gyros.

Smoothing circuit 35 is a second order smoothing cir cuit wherein the second integration is performed by the traverse and elevation gyros on the antenna mount of the launching platform. In actuality, the outputs of the first stage of the smoothing circuit 35 are transmitted to the guidance transmitter for presetting the elevation and traverse gyros of the antenna mount. The gyro pickoifs themselves provide the final outputs of the smoothing circuit and these outputs are used to control antenna power servos thus producing the desired motion of the beam.

In order to provide a cogent detailed description of FIG. it is necessary to start with the general case, where the type A program can be expressed analytically as follows:

0 0 =angular separation between the guidance and tracking beams,

RRm=missile-target range difference, and

the symbol 0 is used to denote conditions at the start of the beam program.

The type A beam program as may be seen by reference to FIG. 3 is a linear function of the difference between the slant ranges to the target and the missile. Target intercept occurs when the range difference reduces to zero. At this point, the angular separation between the guidance and tracking beams also reduces to zero, thereby resulting in a crossover between the two beams. The type B program is defined by the cosine-squared expression:

FIG. 4 illustrates the type B program as having a shape that approaches zero asymptotically in the vicinity of the target intercept point. This makes an exact measure of the time-to-intercept unnecessary and thus the type B program depends less upon range than does the type A program.

In generating the type B beam program, the computer utilizes the properties of a second-order smoothing circuit shown in FIG. 5. The first stage of integration is provided by the computer and the second by the guidance radar closed loop response. Target and beam positional data are compared in differential 71 to produce an error.

6 This r'rtn, e, is multiplied in multipliers 72 and 73 by the beam sensitivity parameters 2wn and (wn 2m) respectively. The output from multiplier 73 is combined with the output from multiplier 72 in differential 76 after it has been integrated in integrator 74 to generate a guidance beam rate. This rate is then converted into an angular output by the guidance radar servo system represented symbolically by integrator 77. Depending on whether the type A or type B program is considered, the type of operation of the guidance control loop is either predominantly steady-state or predominantly transient. The type A program results in steady-state operation, whereas the type B program utilizes the transient response of the guidance loop.

When the switch in FIG. 5 is in the B position, the following parametric equations can be written:

The equation which relates the beam separation angle to the target in the beam program is the one of interest. Eliminating 0 and V from the above three equations yields the desired equation:

Equation 4 can be used to demonstrate that in the steady-state condition, the circuit does not exhibit a lag due to input rates. Assuming that on is a constant, the Laplace transform of this equation is given by:

A series expansion of the foregoing expression in increasing powers of s results in the following error expression:

Since the lowest order term of Equation 4 is represented by the second derivative (or acceleration) of the input, it is apparent that the steady-state error is zero for a nonaccelerating input.

Considering the physical properties of the circuit of FIG. 5 also makes it apparent that the circuit does not have rate lag. For a constant input rate, the first integrator charges up to the desired rate as the error in the loop reduces to zero. When the steady-state condition is reached, the stored rate in this integrator maintains a constant rate output and the loop operates with zero error. Constrasting this circuit with first order circuits with respect to system lags, this circuit has an advantage over a first order circuit because a comparable degree of accuracy is obtainable with a lower level of loop sensitivity. Used in conjunction with the beam programmer of the present invention, this feature is desirable because a decrease is required loop gain results in decreased system bandwidth and, consequently, in heavier smoothing. This is desirable since it is only necessary to provide sufiicient gain near target intercept to enable the system to respond to target maneuvers.

In the generation of the type B beam program proper control of the beam sensitivity parameters makes it possible to match the step-function response of the guidance loop to the cosin-squared program shape. To obtain the required function for this control it is necessary to sub stitute the appropriate time expressions determined from the cosine-squared formula for e and its derivatives in Equation 4. From this a function of on is obtained which, if it is applied to the guidance loop, controls the error response to produce a cosin-squared program shape. In practice, it is not necessary to simulate the complex function exactly and a sufficiently close approximation to the desired curve may be obtained by allowing wit to vary inversely with time-to-intercept by the following equation:

Substituting this equation into the Equation 4 results in the following expression of Equation 4:

The solution of the above equation represents the beam program and is given by the expression (where at Tt=-Tpt, e=e and =)2 For the type A beam program, the guidance radar must respond to keep the error between the forcing function, K(RR in FIG. and the beam separation angle small. To accomplish this, the guidance loop should approach steady-state operation as soon as practicable. This requirement for a relatively rapid response calls for loop sensitivities which are substantially higher than those normally encountered in the type B program. While additional loop gain is supplied to increase the response, caution has been exercised to guard against too great a loss in system smoothing. Therefore, an arrangement which maintains wn constant during a major portion of the trajectory is required. Since increased system sensitivity necessary to insure proper target intercept is required only near the end of the beam program, wn starts out being a constant and, at a predetermined time before intercept, switches over to a variable mode. The present computer arrangement features an wn='U.16 radian per second until 52.5 seconds before intercept. Below this value of time-tointercept, wn increases to 1 radian per second according to the following relationship:

Tt+Kb 5) This arrangement produces three categories of system response:

(a) settling time fixed,

(b) settling time a given percentage of total time-of flight, and

(c) a combination of both.

In the case of a stationary target and assuming that the range to the missile varies linearly, t e above equation reduces to:

This equation is similar in form to the error equation for the type B program but e is now representative of the error between the forcing function and the missile-target angular difference. Initially, the beam separation angle is equal to the forcing function which makes 5 equal to zero. Since the system is initially at rest, is equal to the rate of change of the forcing function or:

Where the above initial conditions are present and for g equal to unity and on equal to a constant, the solution to the second equation above is:

When:

Ka Tt+Kb rn /4Ka+1 2 2 where:

+2;Ka+t H FFT) and:

In the discussion on beam guidance presented above which described the operation of the guidance loop it was assumed that it was for a shipboard system in which all deck-referenced information had been transformed into stable quantities. Actual control of the guidance beam would then be preformed in the vertical and horizontal planes with the beam being programmed separately in elevation and bearing. However, such a system requires several major coordinate transformations with each involving a long chain of resolvers. As conventional resolvers do not possess sufiicient accuracy to meet the collimation and stabilization requirements of the system, this procedure for generating the beam program is not practicable. Therefore, a concept whereby the motion of the guidance beam is controlled at the plane determined by the guidance beam itself and the line-of-sight to the target was adopted. With the guidance radar control axes selected as the basic reference for the system, the guidance picture is one in which vectors are used to represent the relative position of the guidance beam with respect to the target line of sight. In order to facilitate a discussion of this vector concept, a brief description of the operation of guidance radar operation will be given.

As pointed out above, the second stage of integration in the guidance loop is provided by the guidance radar. This is due to the fact that the guidance radar behaves as a rate servo in its response to orders from the computer. FIG. 6 illustrates a simplified symbolic block diagram of the guidance transmitter servo loop. Rate orders from the computer cause a physical torque to be applied on the gimbal of gyro '61. Gyro in turn generates a proportional voltage which drives motor 64 after passing through preamplifier 62 and amplifier 63. The motor rotates the antenna mount about the axis on which gyro 61 is mounted. The output of motor 64 is fed to differential 66 the second input of which is s-hips motion. This differential is representative/of the transmission of ships motion to the antenna line of sight through the antenna mount. Rate sensor 67 which is symbolic of the rate sensing action of the gyro is connected between differential 66 and gyro 61 and counteracts the torque on the gimbal of gyro 61. When the torque applied by rate sensor 67 to gyro 61 has matched the applied torque caused by the input from the computer the guidance mount has assumed the rate called for by gyro rate orders.

A beam coordinate system is used as the basic reference for the guidance loop of the guidance radar with the coordinate system being represented by a set of orthogonal unit vectors along the elevation, beam, and traverse axes of the antenna mount. A rotation about the line-of-sight occurs when ships motion is present. This rotation about the line-of-sight gives rise to several quantities upon which the computer is dependent for successful operation, namely, the cross-traverse angle and the cross-traverse rate. The significance of these quantities will become apparent in the following description of the guidance concept as it applies to the computer.

FIG. 7 illustrates the vector relationship between the guidance and tracking beams. As shown in FIG. 7 a boresight datum plane has been constructed perpendicular to the tip of the line of sight or 2 axis of the beam coordinate system. The vector A lies along gyro reference axis and terminates at the boresight plane. This vector represents the desired line-of-sight of the guidance beam as determined by the torques exerted on the gyros by the computer. The vector 2 represents the actual direction of the beam and is ordinarily coincident with A. Any separation between the two constitutes a lag in the power servos of the guidance mount. M is a vector also terminating at the boresight plane. However, this vector points in the direction of the target. The space separation between the two beams is given by (A If), which is the vector in the boresight plane connecting the tips of the g and M vectors. The separation in the boresight plane between the programmed and actual positions of the guidance transmitter is denoted by (5-2). The separation between the programmed vector and the AI vector is represented by the vector A, or (l[ Proper control of A in magnitude and direction can program the A vector toward the target line-of-sight. Therefore, the A vector is directly analogous to the quantity 6 in Equation 3a. Also, the vector rates and accelerations are analogous to the angular rates and accelerations given in Equations 3b and 30. Thus, for the type B program, the vector solution to the problem is given by the following equations:

where 1 (2 I) g represents the component of perpendicular to g, A i+ (24) 2 represents the component of A perpendicular to g and K-=O.

When the guidance loop is mechanized to solve the above equations, A reduces to zero in accordance with the type B beam program. The shape of the program is slightly distorter for cases in which the beam separation is large. However, the program distortion is a minor consideration because the dynamic properties of the beam program are not altered.

In order to utilize equations 7a, 7b and 7c it is necessary to express them in terms of components along the elevation and traverse axes of the guidance radar as follows:

The forcing function must be converted into vector form as illustrated in FIG. 8 for the type A beam program. A vector C has been added such that A is now given by Q-. Also, a pseudo-stable, or cross traverse stabilized set of axes is shown, together with the beam coordinates. The essential feature of this pesudo-stable coordinate system is that it carries with it the initial cross-traverse 10 orientation of the guidance beam for each particular beam program.

In the generation of the type A beam program, the forcing function is first referenced to this coordinate system and then transformed into beam coordinates by a resoltuion through the incremental change in the crosstraverse angle. The vector Q moves in accordance with the forcing function and the vector A supplies the error which is used to enable the vector A to keep up with Q. The equation for A is as follows:

where Z and 2 represent, respectively, the local lateral and vertical references axes in the boresight plane, and f and g are magnitudes of the respective components of the foreing function along the pseudo-stable reference axes. The f and g scalar quantities determine the shape of the beam program and are given as follows:

The instantaneous cross-traverse orientation of the guidance transmitter is related to its original orientation by the following relationship:

7:; cos (AZms) sin (AZms) 2:; sin (AZms) cos (AZms) When the above equations are combined there results the following scalar control equations:

The remaining operations in the guidance loop are identical to that of the type B program.

The traverse and elevation components of the vector separation (y-g) FIG. 7 between the two beams is established by the input computations. The component development is based on a knowledge of the deck-referenced angles of the tracking and guidance radars. The coordinate transformation required for this development appears in FIG. 9. The unit vector I along the target lineof slight has components r r and r along the three axes of the guidance transmitter coordinate system. The M vector is an extension of the unit vector out to the boresight plane. The component of l\ l along the unit vector g is equal to unity by definition and it follows from the geometry that:

Referring now more particularly to FIG. 10 there is shown an embodiment of the present invention comprising five basic elements: input computer 31, function generator 33, beam sensitivity computer 32, cross-traverse rate circuit 34, and smoothing circuit 35. The embodiment shown in FIG. 10 is analogous to the second order smoothing circuit of FIG. 5 with the exception that FIG. 5 shows second stage integration. However, as previously pointed out second stage integration is provided by the guidance radar closed loop response illustrated in FIG;

6. Therefore, the output of summation circuits 51 and 54 are first stage integration quantities which are used to continuously position the guidance beam of the guidance radar wherein the second integration is inherently performed. For the type B trajectory switches S and S; are in the position shown in FIG. 10. In other Words, function generator 33 is disconnected from smoothing circuit 35.

Input computer 31 provides smoothing circuit 35 with the traverse and elevation components of the vector separation gg between the two beams, that is, between the actual line-of-sight of the radar guidance beam and the target line-of-sight. The traverse component M of the vector separation is fed to summation network 36 while the elevation component M of the vector separation is fed to summation network 37 of smoothing circuit 35 as well as to function generator 33 discussed hereinbelow. The traverse component A and the elevation component A of the separation in the boresight plane (as discussed in reference to FIGS. 7 and 8) between the programmed and actual positions of the guidance transmitter are respectively fed into summation circuits 36 and 37 of smoothing circuit 35. Summation circuits provide the traverse and elevation separation quantities, A and A as inputs to multiplier circuits 39 and 49, respectively. A and A are the scalar components of A, the tangent of the space angle between the target and missile lines of sights as shown in FIG. 7. The guidance radar servo 40 feeds back to summation circuits 37 and 36 inputs proportional to the guidance elevation error A and guidance traverse error A respectively.

The beam sensitivity parameters 2wn and (ton -2 min) are used to control the rate of closure of the beam separation angle in a manner readily seen in the preceding mathematics. These parameters are computed in the beam sensitivity computer 32 in the following manner. Divider circuit 41 receives inputs from the guidance radar 50 proponional to missile-target range difference rate and another input proportional to missile-target range difference which are divided to provide an output proportional to time-to-intercept. This output is added to a constant Kb in summation network 42 and is provided as an input to divider circuit 43. Divider circuit 43 is provided with a second input Ka which is a nondimensional constant. The angular output of divider circuit 43 on of the inverse of time-to-intercept is applied as an input to multiplier circuit 44 and squaring circuit 45. The second input to multiplier circuit 44 is the damping constant 2; and the output from multiplier circuit 44 is one of the desired quantities 25am which is applied as an input to multiplier circuit 38 of smoothing circuit 35.

The output from squaring circuit 45 is fed as an input to multiplier circuit 46. The nondimensional constant Ka which is fed as an input to divider circuit 43 is changedinto a function of g or l2/Ka by means of function generator 47 and applied as the second input to multiplier circuit 46. The output 9i Ito which is equivalent to the desired quantity wn -2wn is provided as an input to multiplier circuit 39 of smoothing circuit 35.

The output from multiplier circuit 49 of beam sensitivity circuit 32 is also fed as an input to multiplier circuit 48 of smoothing circuit 35. The output from multiplier circuit 46 of beam sensitivity circuit 32 is also fed as an input to multiplier circuit 49 of smoothing circuit 35. Multiplier circuits 38 and 48 receive inputs from summation circuits 36 and 37, respectively, and therefore have outputs proportional to 2wnAl and Zg'wnAB, respectively.

The output of multiplier circuit 38 is fed as one input to summation network 51 while the output from multi- 12 plie'r 39 is fed as one input to the summation network 52. The output from multiplier circuit 49 is fed as one input to summation circuit 53 while the output from multiplier circuit 48 is fed as one input to summation circuit 54.

A measure'of the cross-traverse rate (angular rate about the guidance beam axis) is required to compensate for con'olis accelerations and to compute the incremental cross-traverse angle AZms for the type A beam program. A computation of AZms requires extremely accurate rate data and for this reason cross-traverse rate circuit 34 which utilizes a high gain second order loop around the cross-traverse gyro is required. The cross-traverse rate circuit 34 also computes the correction terms required for the gyro cross feed in the smoothing circuit 35. The correction terms appear in Equations 8a and 8 Therefore, cross-traverse rate circuit 34 provides the measure of the input (DZms)s or 40 as an input to each of multipliers 56 and 57 as well as to function generator 33. The input from summation circuit 52 is integrated in integrator 58, the output of which provides one input to multiplier 57. The output of summation circuit 53 is integrated in integrator circuit 59, the output of which serves as one input to multiplier circuit 56. The outputs of integrator circuits 58, 59 are also connected to summation circuits 51 and 54, respectively. The output of summation circuit 51 is the traverse gyro order which is fed to the gyro of the antenna mount of the guidance radar. The output from summation circuit 54 is the elevation gyro order which is fed to the gyro of the guidance radar. The outputs from summation circuits 51 and 54, respectively, position the guidance beam in the proper traverse and elevation poitions.

Function generator 33 provides the elevation and traverse components of the forcing function in the guidance beam coordinate system for the type A program. As previously explained no forcing function is required in the type B program. This operation consists of storing and locking, at the time of capture, the elevation and traverse components of the (g A) vector shown in FIG. 8 and the cross-traverse orientation of the guidance beam. The locked elevation and traverse components are varied in accordance with the type A closure by means of the input representative of range difference between missile and target and resolved through a change in cross-traverse to produe the desired output for the smoothing circuit. The incremental change in the roll angle AZms of the guidance antenna is obtained by direct integration of the cross-traverse rate (Dzm's)s.

In other words, the function generator 33 when switches S1 and S2 are in the position opposite from that shown inserts into summation circuits 36 and 37 and the additional desired traverse separation and the de sired elevation separation, respectively, necessary to provide in the system the up and over trajectory of type A. In either case, the guidance beam from the guidance radar is positioned by the gyros to reduced separation angle as a function of change in range between the missile and target which range becomes zero when the tracking and guidance beam coincide in three dimensional coordinate system.

The outputs from summation circuits 51 and 54 are proportional to the designated true target bearing and guidance elevation gyro order, respectively. These outputs up to but before insertion into the guidance radar servo 40 are the solutions of Equations 82 and 8 for the general case where m (angular rate about the elevation axis) and m (cross-traverse rate about the traverse axis) are not equal to zero. Upon insertion into the guidance radar servo 40 the second integration is performed upon the quantities expressed in Equations 8e and 8 making the solution complete.

Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within 13 the scope of the appended claims, the invention may be practiced otherwise than as specifically described.

What is claimed is: 1. A computer system for programming the flight of a beam riding missile subsequent to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

computer circuit means, first means providing said computer circuit means with inputs representative of the traverse and elevation components of the vector separation between the actual direction of the guidance beam and thedesired direction ofthe guidance beam, second means providing said computer circuit means with inputs representative of the traverse and eleva tion components of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter, beam sensitivity computer means providing inputs representative of beam sensitivity parameters to said computer circuit for controlling the rate of closure of the angle separating actual and desired guidance beams as a function of the remaining time of flight of the missile, cross-traverse rate circuit means connected to said computer circuit means providing said computer circuit means with an input representative of the crosstraverse rate about the beam axis, guidance radar servo means connected to the output of said computer circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders for positioning the guidance transmitter in accordance with the beam program. 2.'A computer system according to claim 1 wherein said beam sensitivity computer comprises:

divider circuit means having inputs representative of range between a missile and target and rate of change of said range between a missile and target having an output representative of time-to-intercept between the missile and target, first circuit meansconne cted to said divider circuit for converting said output representative of time-tointercept into a linear angular function of said timeto-intercept, second circuit means connected to said first circuit means converting said linear angular function into a nonlinear angular function of said time-to-intercept. 3. A computer system for programming the flight of a beam riding missile subsequent to lauch by positioning the guidance beam of a guidance transmitter, comprising in combination:

first summation circuit means, second summation circuit means, first means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the actual direction of'the guidance beam and the desired direction of the guidance beam and to said second summation circuit means with an input representative of the elevation component of the vector separation between the-actual direction of the guidance beam and the desired direction of the guidance beam, second means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter and an input to said second summation circuit means and providing said second summation circuit means with an input representative of the elevation component of .the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter, first computer means connected to said first summation circuit means,

second computer means connected to said second summation circuit means, beam sensitivity computer means providing an input to said first and second computer means of the angular function of the remaining time of flight 0f the missile, cross-traverse rate circuit means connected to said first and second computer means providing said first and second computer means with inputs proportional to the cross-traverse rate about the beam axis. guidance radar servo means connected to the output of said first and second computer circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders from said first and second computer means respectively for positioning the guidance transmitter in accordance with beam program. 4. A computer system according to claim 3 wherein said beam sensitivity computer comprises:

divider circuit means having inputs representative of range between a missile and target and rate of change of said range between a missile and target having an output representative of time-to-inter-cept between the missile and target, first circuit means connected to said divider circuit for converting said output representative of time-tointercept into a linear angular function of said timeto-intercept, second circuit means connected to said first circuit means converting said linear angular function into a nonlinear angular function of said time-to-intercept. 5. A computer system for programming the flight of a beam riding missile subsequent to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

first summation circuit means, second summation circuit means, input computer means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam and said second summation circuit means with an input representative of the elevation component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam, guidance radar servo means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter and providing said second summation circuit means with an input representative of the elevation component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter, first, second, third and fourth multiplier circuit means, means connecting the output of said first summation circuit means to said first and second multiplier circuit means, means connecting the output of said second summation circuit means to said third and fourth multiplier circuit means, beam sensitivity computer means providing an input representative of a first beam sensitivity parameter to said first and fourth multiplier circuit means and an input representative of a second beam sensitivity parameter to said second and third multiplier circuit means for controlling the rate of closure of the angle separating actual and desired guidance beam as a function of the remaining time of flight of the missile, first computer means connected to said second multiplier circuit means,

second computer means connected to said third multiplier circuit means,

third summation circuit means connected to said first multiplier circuit means and said first computer means,

fourth summation circuit means connected to said fourth multiplier circuit means and said second computer means, cross-traverse rate circuit means connected to said first and second computer means and said third and fourth summation circuit means providing said first and second computer means with an input proportional to the cross-traverse rate about the beam axis,

said cross-traverse rate circuit providing said third summation circuit means with an input representative of said cross-traverse rate multiplied by the traverse component of the vector separation between the programmed and actual positions of the guidance transmitter and said fourth summation circuit means with an input representative of said crosstraverse rate multiplied by the elevation component of the vector separation between the programmed and actual position of the guidance transmitter,

guidance radar servo means connected to the output of said third and fourth summation circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders for positioning the guidance transmitter in accordance with the beam program.

6. A computer system according to claim wherein said beam sensitivity computer comprises:

divided circuit means having inputs representative of range between a missile and target and rate of change of said range between a missile and target having an output representative of time-to-intercept between the missile and target,

first circuit means connected to said divider circuit for converting said output representative of time-tointercept into a linear angular function of said timeto-intercept,

second circuit means connected to said first circuit means converting said linear angular function into a nonlinear angular function of said time-tointercept. 7. A computer system for programming the flight of a beam riding missile subsequent to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

first summation circuit means, second summation circuit means, first means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam and said second summation circuit means with an input representative of the elevation component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam, second means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter and an input to said second summation circuit means and providing said second summation circuit means with an input representative of the elevation component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter,

first, second, third and fourth multiplier circuit means,

means connecting the output of said first summation circuit means to said first and second multiplier circuit means,

16 means connecting the output of said second summation circuit means to said third and fourth multiplier circuit means, I I I II I beam sensitivity computer means providing an input 5 representative of a first beam sensitivity parameter to said first and fourth multiplier circuit means and an input representative of a second beam sensitivity parameter to said second and third multiplier circuit means for controlling the rate Qf closure of the angle separating actual and desired guidance beams as a function of the remaining time of flight of the missile, I I I third summation circuit means connected to said second multiplier circuit means, I II fourth summation circuit means connected to said first multiplier circuit means, I fifth summation circuit means connected to said third multiplier circuit means, I I I sixth summation circuit means connected to said fourth multiplier circuit means, I I fifth multiplier circuit means having an output connected to said third summation circuit means, sixth multiplier circuit means having anoutput connected to said fifth summation circuit means, cross-traverse rate circuit means connected to said fifth and sixth multiplier circuit means and providing'said fifth and sixth multiplier circuit means with an input proportional to the cross-traverse rate about the beam axis, I I I I I said cross-traverse rate circuit providing said fourth summation circuit means with an input representative of a first modified cross-traverse rate and said sixth summation circuit means with an input repre- I sentative of a second modified cross-traverse rate, first integrator circuit means having an input connected to said third summation circuit means and an out-put connected to, said fourth summation circuit means and said sixth multipliercircuit means, second integrator circuit means having its input connected to said fifth summation circuit means and its output connected tp sai d sixth summationcircuit means and said fifth multiplier circuit means, guidance radar servo means connected to the output of said fourth and sixth summation circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders for positioning the guidance transmitter in accordance with the beam program. I I I I II 8. A computer system according to claim 7 wherein said beam sensitivity computer comprises: I I I divider circuit means having inputs representative of range between a missile and target and rate of change of said range between a missile and target having an output representative of time-to-intercept between, the missile and target, I I I I first circuit means connected to said divider circuit for converting said output representative of I time-tointercept into a linear angular function of said timeto-intercept,- I I v I second circuit means connected to said first circuit means converting said linear angular function into a nonlinear angular function of said time-to-intercept. 9. A computer system for programming the flight of a beam riding missile in one of two modes subsequent to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

computer circuit means, function generator means for developing traverse and elevation components of desired angle separation between the radar tracking and guidance beams as a function of range, I switch means for selectively coupling said function generator means to provide said computer circuit means with said traverse and elevation components, respectively, as inputs thereto,

17 first means providing said computer circuit means with inputs representative of the traverse and elevation components of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam, second means providing said computer circuit means with inputs representative of the traverse and elevation components of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter,

beam sensitivity computer means providing inputs representative beam sensitivity parameters to said computer circuit for controlling the rate of closure of the angle separating actual and desired guidance beams as a function of the remaining time of flight of the missile,

cross-traverse rate circuit means connected to said computer circuit means providing said computer circuit means with an input representative of the crosstraverse rate about the beam axis,

guidance radar servo means connected to the output of said computer circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders for positioning the guidance transmitter in accordance with the beam program.

10. A computer system according to claim 9 wherein said beam sensitivity computer means comprises:

divider circuit means having inputs representative of range between a missile and target and rate of change of said range between a missile and target having an output representative of time-to-intercept between the missile and target,

first circuit means connected to said divider circuit for converting said output representative of time-to-interccpt into a linear angular function of said time-tointercept,

second circuit means connected to said first circuit means converting said linear angular function into a nonlinear angular function of said time-to-intercept.

11. A computer system for programming the flight of a beam riding missile in one of two modes subsequent to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

first summation circuit means,

second summation circuit means,

function generator means for developing traverse and elevation components of desired angle separation between the radar tracking and guidance beams as a function of range,

switch means for selectively coupling said function generator means to provide said first and second summation circuit means with said traverse and elevation components, respectively, as inputs thereto, first means providing said first summation circuit means with an input representative of the traverse compo nent of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam and to said second summation circuit means with an input representative of the elevation component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam,

second means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter and an input to said second summation circuit means and providing said second summation circuit means with an input representative of the elevation component of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter,

'first computer means connected to said first summation circuit means, second computer means connected to said second summation circuit means,

beam sensitivity computer means providing an input to said first and second computer means of the angular function of the remaining time of flight of the missile,

cross-traverse rate circuit means connectedto said first and second computer means providing said first and second computer means with inputs proportional to the cross-traverse rate about the beam axis,

guidance radar servo means connected to the output of said first and second computer circuit means whereby said guidance radar servo means is provided with traverse and elevation gyro orders from said first and second computer means respectively for positioning the guidance transmitter in accordance with the beam program. 12. A computer system for programming the flight of a beam riding missile in one of two modes subsequent to launch by positioning the guidance beam of .a guidance transmitter, comprising in combination:

first summation circuit means, second summation circuit means, function generator means for developing traverse and elevation components of desired angle separation between the radar tracking and guidance beams as a function of range, switch means for selectively coupling said function generator means to provide said first and second summation circuit means with said traverse and elevation components, respectively, as inputs thereto, input computer means providing said first summation circuit means with an input representative of the traverse component of the vector separation between the the actual direction of the guidance beam and the desired direction of the guidance beam and said second summation circuit means representative of the elevation component of the vector separation between the actual direction of the guidance beam and the desired direction of the guidance beam,

guidance radar servo means providing said first summation circuit means with an input representative of the traversecomponent of the vector separation between the programmed position of the guidance transmitter and the actual position of the guidance transmitter and an input to said second summation circuit means and providing said second summation circuit means with an input representative of the elevation component of the vector separation between the' programmed position of the vguidance transmitter and the actual position of the guidance transmitter, first, second, third and fourth mulitplier circuit means, means connecting the output of said first summation circuit means to said first and second multiplier circuit means,

means connecting the output of said second summation circuit means to said third and fourth multiplier circuit means, beam sensitivity computer means providing an input representative of a first beam sensitivity parameter to said first and fourth multiplier circuit means and an input representative of a second beam sensitivity parameter to said second and third multiplier circuit means for controlling the rate of closure of the angle separating actual and desired guidance beam as a function of the remaining time of flight of the missile,

first computer means connected to said second multiplier circuit means,

second computer means connected to said third multiplier circuit means,

19 third summation circuit means connected to said first multiplier circuit means and said first computer means, fourth summation circuit means connected to said cuit means to said first and second multiplier circuit means,

means connecting the output of said second summation circuit means to said third and fourth multiplier cirfourth multiplier circuit means and said second comcuit means,

puter means, beam sensitivity computer means providing an input cross-traverse rate circuit means connected to said first representative of a first beam sensitivity parameter and second computer means and said third and to said first and fourth multiplier circuit means and fourth summation circuit means providing said first an input representative of a second beam sensitivity and second computer means with an input proporparameter to said second and third multiplier circuit tional to the cross-traverse rate about the beam axis, means for controlling the rate of closure of the ansaid cross-traverse rate circuit providing said third sumgle separating actual and desired guidance beams as mation circuit means with an input representative of a function of the remaining time of flight of the missaid cross-traverse rate multiplied by the traverse sile,

component of the vector separation between the prothird summation circuit means connected to said secgrammed and actual positions of the guidance transond multiplier circuit means,

mitter and said fourth summation circuit means with fourth summation circuit means connected to said first an input representative of said cross-traverse rate multiplier circuit means,

multiplied by the elevation component of the vector fifth summation circuit means connected to said third separation between the programmed and actual posimultiplier circuit means,

tion of the guidance transmitter, sixth summation circuit means connected to said fourth guidance radar servo means connected to the output multiplier circuit means,

of said third and fourth summation circuit means fifth multiplier circuit means having an output conwhereby said guidance radar servo means is pronected to said third summation circuit means,

vided with traverse and elevation gyro orders for sixth multiplier circuit means having an output conpositioning the guidance transmitter in accordance with the beam program. 13. A computer system for programming the flight of nected to said fifth summation circuit means, cross-traverse rate circuit means connected to said fifth and sixth multiplier circuit means and providing said a beam riding missile in one of two modes subsequent'to launch by positioning the guidance beam of a guidance transmitter, comprising in combination:

fifth and sixth multiplier circuit means with an input proportional to the cross-traverse rate about the first summation circuit means, second summation circuit means, function generator means for developing traverse and beam axis,

said cross-traverse rate circuit providing said fourth summation circuit means with an input representative of a first modified cross-traverse rate and said elevation components of desired angle separation besixth summation circuit means with an input repretween the radar tracking and guidance beams as a sentative of a second modified cross-traverse rate, function of range, first integrator circuit means having an input connected switch means for selectively coupling said function gento said third summation circuit means and an output erator means to provide said first and second sumconnected to said fourth-summation circuit means mation circuit means with said traverse and elevation and said sixth multiplier circuit means, components, respectively, as inputs thereto, second integrator circuit means having its input confirst means providing said first summation circuit means nected to said fifth summation circuit means and its with an input representative of the traverse compooutput connected to said sixth summation circuit nent of the vector separation between the actual dimeans and said fifth multiplier circuit means, rection of the guidance beam and the desired direcguidance radar servo means connected to the output of tion of the guidance beam and said second summasaid fourth and sixth summation circuit means wheretion circuit means with an input representative of the -by said guidance radar servo means is provided with elevation component of the vector separation betraverse and elevation gyro orders for positioning the tween the actual direction of the guidance beam and g idance transmitter in accordance with the beam the desired direction of the guidance beam, program. second means providing said first summation circuit References Cited means with an input representative of the traverse UNITED STATES PATENTS component of the vector separation between the programmed position of the guidance transmitter and 3169'727 2/1965 Schmader et 244 14 the actual position of the guidance transmitter and 3,164,339 1/ 1965 Schroeder et a1. 244-14 an input to said second summation circuit means and 3,008,668 11/1961 Darlington 244-14 providing said second summation CII'CLllt means with 2,745,095 5 1956 Stoddard 244 14 an input representative of the elevation component of the vector separation between the programmed 0 RICHARD A. FARLEY, Primary Examiner.

BENJAMIN A. BORCHELT, Examiner.

M. F. HUBLER, Assistant Examiner,

position of the guidance transmitter and the actual position of the guidance transmitter, first, second, third and fourth multiplier circuit means, means connecting the output of said first summation cir- 

1. A COMPUTER SYSTEM FOR PROGRAMMING THE FLIGHT OF A BEAM RIDING MISSILE SUBSEQUENT TO LAUNCH BY POSITIONING THE GUIDANCE BEAM OF A GUIDANCE TRANSMITTER, COMPRISING IN COMBINATION: COMPUTER CIRCUIT MEANS, FIRST MEANS PROVIDING SAID COMPUTER CIRCUIT MEANS WITH INPUTS REPRESENTATIVE OF THE TRAVERSE AND ELEVATION COMPONENTS OF THE VECTOR SEPARATION BETWEEN THE ACTUAL DIRECTION OF THE GUIDANCE BEAM AND THE DESIRED DIRECTION OF THE GUIDANCE BEAM, SECOND MEANS PROVIDING SAID COMPUTER CIRCUIT MEANS WITH INPUTS REPRESENTATIVE OF THE TRAVERSE AND ELEVATION COMPONENTS OF THE VECTOR SEPARATION BETWEEN THE PROGRAMMED POSITION OF THE GUIDANCE TRANSMITTER AND THE ACTUAL POSITION OF THE GUIDANCE TRANSMITTER, BEAM SENSITIVITY COMPUTER MEANS PROVIDING INPUTS REPRESENTATIVE OF BEAM SENSITIVITY PARAMETERS TO SAID COMPUTER CIRCUIT FOR CONTROLLING THE RATE OF CLOSURE OF THE ANGLE SEPARATING ACTUAL AND DESIRED GUIDANCE BEAM AS A FUNCTION OF THE REMAINING TIME OF FLIGHT OF THE MISSILE, CROSS-TRAVERSE RATE CIRCUIT MEANS CONNECTED TO SAID COMPUTER CIRCUIT MEANS PROVIDING SAID COMPUTER CIRCUIT MEANS WITH AN INPUT REPRESENTATIVE OF THE CROSSTRAVERSE RATE ABOUT THE BEAM AXIS, GUIDANCE RADAR SERVO MEANS CONNECTED TO THE OUTPUT OF SAID COMPUTER CIRCUIT MEANS WHEREBY SAID GUIDANCE RADAR SERVO MEANS IS PROVIDED WITH TRAVERSE AND ELEVATION GYRO ORDERS FOR POSITIONING THE GUIDANCE TRANSMITTER IN ACCORDANCE WITH THE BEAM PROGRAM. 